The present invention is directed to solar power systems for spacecraft, and in particular, improved, lightweight high efficiency solar energy systems for powering spacecraft during long missions in space.
In order to operate during long missions in space, spacecraft generally use solar cell panels to collect solar irradiance and convert it into electrical power to operate the spacecraft attitude control systems, communications, and payloads. Typical spacecraft in geosynchronous orbits (36,800 Km. altitude above the earth surface) require continuous solar power of 3,000 to 60,000 watts. Typical existing solar panels use silicon photocells arranged in two dimensional arrays on a lightweight honeycomb backing. These panels fold up accordion style during launch and unfold once in geosynchronous orbit, where the spacecraft remains above a predesignated geographical point on the earth. At geosynchronous orbit, the earth subtends only 17.5 degrees, and therefore blocks the sun for only 1.16 hours maximum during the Spring and Fall Equinoxes, and not at all during the solstices.
Silicon photocells used without optical concentration provide efficiencies of 12 to 15%, and degrade in output at increased temperatures and as a result of exposure to large quantities of Van Allen radiation. Gallium Arsenide Phosphide (GaAsP) solar cells provide higher efficiencies, but operate best when used in a solar concentrator.
At their present high state of development, solar panels for spacecraft are evaluated on so-called "Figures of Merit" such as watts/Kg. or watts/dollar. It is the primary object of the present invention to provide a solar powered system for an orbiting spacecraft exhibiting significant advantages in such "Figures of Merit". This object is achieved by providing a lightweight, high efficiency solar power system for a spacecraft which automatically deploys from its launch to its operating mode, in a very simple and space-efficient manner, as will be discussed herein.